1. Field of the Invention
This invention relates to aircraft laminar flow control (LFC) systems, and more specifically, to a method and system using plasma discharge to encourage laminar flow along the surface of a wing airfoil.
2. Description of the Related Art
Since the 1930s, laminar flow control (LFC) has been touted as the technology that would enable aircraft to sip fuel and shrug off drag, slipping through the air with the greatest of ease without reducing lift. Laminar flow is achieved by reducing the magnitude of disturbances and instabilities in the very thin and relatively stagnant layer of air between the skin of an aircraft and the free-stream air surrounding it call the xe2x80x9cboundary layer.xe2x80x9d By keeping these fluctuations small, the nonlinear interactions leading to turbulence can be curtailed and/or delayed. Currently, the most robust methods for controlling the disturbance amplitudes are based on modifying the boundary layer mean flow via airfoil geometry (i.e., by tailoring the pressure gradient, Cp) or by applying surface suction. However, these methods have not delivered on the promise of LFC.
Since modifications to the pressure gradient do not actively consume power, this approach has been termed xe2x80x9cnatural laminar flowxe2x80x9d. The successful application of this approach and attainment of drag reduction benefits has been demonstrated both theoretically and in testing for nominally two-dimensional boundary layers. The main disadvantage of the natural laminar flow approach is that the modified Cp distribution is generally unacceptable from an overall airplane performance point of view. For this reason, natural laminar flow is not frequently used for increasing the extent of laminar flow.
The use of suction has also been successfully tested to show improved laminar flow and reduced drag without the adverse restrictions on the Cp distribution. However, the suction approach has its own shortcomings, including increased costs, added weight, and increased complexity of the overall flow-control system as compared to the baseline non-suction configuration. These shortcomings partially offset the performance savings. There are also potential performance penalties associated with suction applications, e.g., suction drag and increased roughness sensitivity due to thinner boundary layers. Additionally, the porous suction surface can require increased maintenance.
It is also known to use a combination of suction and pressure gradient tailoring (termed xe2x80x9chybrid laminar flow controlxe2x80x9d) to effectively achieve laminar flow with more practical Cp distributions. While the overall performance of the aircraft is improved to acceptable levels, the hybrid laminar flow control approach still suffers the shortcomings of the suction system.
The application of surface air cooling (to below the adiabatic surface temperature) has also been theorized to be an effective flow control technique. The general theory predicts that cooling of an airflow surface to lower than the adiabatic surface temperature will cool the passing boundary layer, which in turn will slow the development and growth of instabilities. Conceived surface cooling techniques, however, are thought to be impractical for large surface areas such as those in a large commercial transport. Because of this, the idea of surface cooling is not exploited in current aircraft configurations.
The beneficial effects of surface cooling have also been theorized to occur by application of local heat to a stable upstream region of the boundary layer. In theory, the heated upstream boundary layer then encounters a cooler downstream surface to result in a net temperature decrease experienced by the boundary layer that is similar to the net change in temperature achieved by simply cooling the downstream surface. This approach was demonstrated experimentally at TsAG1 and at I.T.A.M. in Russia during the mid-to-late 1980""s. Specifically, the results showed that increased laminar flow could be achieved by localized heating in the leading-edge region of a flat plate. (See for example, Dovgal, A. V., Levchenko, V. Ya. and Timofeev, V. A. (1990) xe2x80x9cBoundary layer control by a local heating of the wall,xe2x80x9d from: IUTAM Laminar-Turbulent Transition, eds. D. Arnal and R. Michel, Springer-Verlag, pp. 113-121). One of the problems in applying this alternative technique to airfoils has been the loss of performance benefit after only a relatively short period of time due to the transfer of heat from the boundary layer flow to the cooler surface downstream. As heat is transferred from the boundary layer flow, the surface temperature rises and the relative temperature difference between the flow and the surface diminishes. This reduces the stabilizing effect on the boundary layer and eventually terminates the laminar-flow benefit.
U.S. Pat. No. 6,027,078 to Crouch provides a localized heating system for use with an airfoil having a leading edge region, a controlled surface extending aft from the leading edge on one side of the airfoil, and an uncontrolled surface extending aft from the leading edge region on the opposite side of the airfoil. The heating system includes an electro-thermal heat source located at the leading edge region; and a heat sink positioned aft of the heat source and adapted for heat transfer from the controlled surface to the uncontrolled surface. The electro-thermal heat source modifies the pressure distributions around the airfoil to improve laminar flow. The transfer of heat by the heat sink improves the boundary layer airflow along the controlled surface. Crouch""s electro-thermal surface heating has a long relaxation time that makes it ineffective in responding to real-time changes in flight conditions and is very inefficient in heating flow over the vehicle surface.
In summary, the drag reduction benefits of having laminar airflow have been known for many years, however, there are few economically viable laminar airflow control systems available. The general problem has been that the increased costs required to achieve sustained laminar flow substantially erodes the potential benefits. Usually, the laminar flow control system does improve laminar flow over an aerodynamic surface (e.g., wing, nacelle, vertical tail, etc.) and improve overall aircraft performance, but the benefits of the system are more than offset by the increased costs in manufacturing, maintenance, aircraft weight, design complexity, operational costs, reliability, etc. Thus, a need exists for a laminar flow control means that is low cost and low maintenance. The ideal system would further have minimum impact on the weight and configuration complexity of the aircraft.
The present invention provides a system and method for controlling leading edge contamination and crossflow instabilities for laminar flow on aircraft airfoils that is light weight, low power, economical and reliable.
This is accomplished with plasma surface discharges that supply volumetric heating of the supersonic boundary layers to control the Poll Reynolds number and the cross flow Reynolds number and delay transition to turbulent flow associated with the leading edge contamination and crossflow instabilities. A closed-loop feedback control system that incorporates these principles includes three primary components: heat-flow sensors, a PID controller, and plasma discharge elements. Heat-flow sensors distributed around the airfoil surface provide root-mean-square (rms) pulsations of the heat flow to the airfoil skin. These data are fed to the PID controller to determine the flow state (laminar or turbulent) and to drive voltage inputs to the plasma discharge elements, which provide the volumetric heating of the boundary layer on a time scale necessary to adapt to changing flight conditions and delay transition to turbulent flow.
These and other features and advantages of the invention will be apparent to those skilled in the art from the following detailed description of preferred embodiments, taken together with the accompanying drawings, in which: